preliminary design study of - nasa cr-66399 volume 111 summary 14 april 1967 preliminary design...

NASA CR-66399

VOLUME 111 SUMMARY

14 APRIL 1967

PRELIMINARY DESIGN STUDY OF

SLAMAST SCOUT-LAUNCHED ADVANCED MATERIALS

A N D STRUCTURES TEST-BED

VOLUME 111 SUMMARY

s CONTRACT NO. NAS 1-7014

CFSTI PRICE(S) $

)i Hard copy (HC) 3- @

I Microfiche (MF) , &' ff 663 Julv 66

Distribution of this report is provided in the interest of information exchange. Responsibility for the contents resides in the author o r organization that prepared it.

G E N E R A L @ E L E C T R I C 8 u R E-ENT R V SVST E MS DEPARTMENT

https://ntrs.nasa.gov/search.jsp?R=19670021906 2018-05-25T16:50:48+00:00Z

NASA CR-66399

VOLUME 111 S U M M A R Y

14 APRIL 1967

PRELIMINARY DESIGN STUDY OF

SLAMAST SCOUT-LAUNCHED ADVANCED MATERIALS

A N D S T R U C T U R E S TEST-BED

VOLUME 111 SUMMARY

CONTRACT NO. NAS 1-7014

Prepared for

National Aeronautics and Space Administration Langley Research Center

Langley Station Hampton, Virginia

G E N E R A L n .$I E L E C T R I C u

RE-ENTRY SYSTEMS DEPARTMENT A DeporfmerM Of T i e Ilissde a?id Space Division

3198 Chestnut Street , Philadelphia 4. Penna.

PRECEDING PAGE BLANK NOT FiLMED.

ABSTRACT

This report presents the results of a feasibility study of a small, lifting re-entry vehicle test-bed capable of being launched by a Scout launch vehicle and capable of being recovered. The purpose of the study was to determine if it was possible to conduct meaningful, sub-scale, thermostrudural experiments involving panels of interest of representative full scale, manned, lifting re-entry vehicles. The study was based upon the H G l O vehicle concept.

iii

Section

....................................... 4.0 INTRODUCTION 1

4.1 EXPERIMENTFEASIBILITY .............................. 1

Introduction. ....................................... 1 Approach to Thermostructural Similitude .................... 2 Approach to Touchdown. ............................... 3 Simulation Philosophy. ................................ 3

Other Testing Alternatives ............................. 9

Test Panel Design and Trajectory Shaping . . . . . . . . . . . . . . . . . . . 6 Other SLAMAST Benefits. .............................. 8

4.2 DESIGN FEASIBILITY.. ................................. 11

Capability.. ....................................... 13 Flexibi l i t y . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 Redundance.. ...................................... 15

4.3 PROGRAM CONSIDERATIONS. ............................ 16

PXECEDING PAGE BLANK NOT, FIUED.

ILLUSTRATIONS a g e Figure

1

2

3

Elements of Structural Design ........................... 20

Margin of Safety as a Function of Time ..................... ~ 21 ~

I 21 ~

Margin of Safety - Critical Failure Modes ................... Flow Chart Demonstrating Procedure for Selecting SLAMAST Trajectory and Panel Design for HL-10 Simulation. ............. 4

22 ~

Comparison of Maximum Convective Heat Flux Rates for HL-10 and SLAMAST Vehicles.. .............................. 5

23 I 6

7

8

9

10

11

12

13

14

15

16

Maximum SLAMAST Stagnation Heat Transfer Rates . . . . . . . . . . . . 24

SLAMAST Integrated Stagnation History versus Velocity .......... 25 I

SLAMAST Flight Mission Profile, ye = 1 Degree DFI1 ........... 26 I

SLAMASTS/C ...................................... 27

SLAMAST Structural Arrangement and Profile ...... W+ ....... 28

Component Arrangement - Plan and Profile Views. ............. 29

31 I Spacecraft Geometry. .................................

Overall Program Schedule.. ............................ 33

34

38

39

I Analysis Schedules. .................................. Subsystem Development Schedule ......................... Integrated Test Schedule. ..............................

TABLES

18 I 19

I

II

SLAMAST Weight and Balance .......................... SLAMAST Preliminary Parts List ........................

vii

Preliminary Design Study

of

SLAMAST

Scout Launch Advanced Materials and Structures-Bed

Volume III - Summary

INTRODUCTION

This volume of the SLAMAST Design Feasibility Report contains an overall summary of the work done and the conclusions reached during the study. Since the study was composed of three major substudies, this summary is organized accordingly. Hence, summaries are presented for (1) Experiment Feasibility, (2) Design Feasibility, and (3) Program Considerations.

EXPERIMENT FEASIBILITY

Introduction - This section of the report describes an approach to thermostructural similitude for lifting re-entry vehicles in general, with specific emphasis on the HL-10 vehicle. Analysis to date shows that an application of margin of safety philosophy combined with a utilization of traditional similarity parameters, test panel designs, and trajectory shaping holds promise of resulting in useful and valid simulation of full scale MLRV thermostructural systems with an economical suhscale vehicle. In addition, it is shown that in the process of flying a thermostructural similitude mission, valuable data can be simultaneously obtained concerning aerodynamics, thermodynamics, heat shield performance and flow transition.

An approach to the simulation process is presented which permits more flexibility than the usual method involving the duplication of a large number of dimensionless parameters related to an idealized thermostructural situation. This approach is discussed in detail along with its application to the HL-10. Also presented is the numerical substantiation that could be generated during the study period, as well as consideration of the limitations of the simulation technique and a comparison of ground test capabilities versus flight test.

It is believed that the approach developed is both sound and workable. The work done during the study period has not yielded any information to the contrary. The great flexibility in environment generation potentially available with the SLAMAST vehicle tends to reinforce the judgment. However, the numerical analyses which would be generated during the span of this study are not extensive enough nor complete enough to permit an unequivocal statement of feasibility. Regardless of the foregoing, i t is important to note that a test bed of the type analyzed is not and should not be limited to thermostructurd testing alone and, as stated earlier, significant testing can be accomplished in the areas of aerodynamic,thermodynamics, heat shield performance and flow transition.

ADsroach t o Thermostructural Similtude, - The design of a structure, whether it be a flight structure, a civil structure, o r any other type of structure, involves three basic elements: (1) the environment to which the structure will be exposed during its useful life, (2) the characteristics of the materials being employed in the design (i. e., the material properties), and (3) the analysis techniques used to predict the structure's performance, behavior, o r response. The main concern here is, of course, flight structures. The primary items which contribute to each of the three elements for flight structures are shown in Figure 1.

The next point to consider in the development of the simulation approach is the consideration of testing. Tests are performed primaril: for the purpose of proving that the structure will not f a i l in its anticipated environ~~ient, o r that it does not have excessive strength in its environment. In other words, it should be neither overdesigned nor underdesigned.

In the development of manned lifting entry vehicles, there are questions regarding environment, material characteristics, and analysis techniques, hence, the need for testing. While the primary purpose of SLAMAST is thermostructural response simulation, it is impossible to completely isolate this from the environmental and material characteristic considerations. The similitude philosophy developed here takes advantage of this fact, and it will be shown that with the use of the SLAMAST flying laboratory, some verification of environment and material characteristics will be obtained as part of the thermostructural similitude.

The specification of structural sizes, thicknesses, stiffnesses, etc. depends on a knowledge of all three of the basic elements. A deficiency in any one of these areas could result in a structure which is incapable of withstanding the environment o r one that is overdesigned and, hence, overweight. Testing with proper instrumentation will reveal these deficiencies, if they indeed exist. Testing, therefore, becomes a job of demonstrating the accuracy with which environments (pressures, temperature distributions, loads) can be predicted, material behavior in the presence of that environment is known, and structural response in the presence of both these can be predicted.

The final design of a structure is specified in te rms of detail drawings of the structure and a structural analysis of the design. This structural analysis includes a consideration of many o r several types of potential failure modes such as:

(1) Buckling (2) Thermal s t ress (3) Mechanical stress (4) Excessive deflection o r strain

2

I 8

These various structural behaviors need investigation in all the possible flight regimes including:

(1) Launch (2) Abort (3) Space flight (4) Re-entry

a. Nominal b. Overshoot C. Undershoot

Approach to touchdown. - Once the structure has been designed, its capability can, and usually is, presented in t e rms of margins of safety. A margin of safety is given for each potential mode of failure. The margin of safety is defined as

Allowable stress (or strain, etc)

(Actual expected stress) x F.S.

The factor of safety is a number which the expected stress (or load o r strain, etc.) is multiplied by to account for variations in material quality and manufacturing vari- ances. There is a factor of safety for yield (usually 1.0) and an ultimate factor of safety (usually 1.4 o r 1.5 for manned vehicles). Design criteria, by definition, essentially says that no yielding shall exist at yield load and no failures (breaking o r buckling) shall exist at ultimate load.

M.S. = - 1

If the margin of safety for each potential mode of failure is plotted as a function of time, a curve similar to the one shown in Figure 2 would result. If the curves correspond to the modes discussed above, it is seen that the minimum margin of safety (the critical mode of failure) or the "weak-link" in the structure is buckling (point 1) at time tl. This is a prediction based on analytical o r empirical techniques. It tells us that based on our knowledge of the environment, based on our knowledge of material properties, and based on our knowledge of structural response, the structure is closer to buckling than to any other mode of failure. However, since the margin is positive, no failure should occur.

Simulation Philosophy. - With these thoughts in mind, a simulation philosophy o r approach was formulated. This approach consists essentially of reproducing, in an experimental panel on a subscale maneuvering vehicle, the minimum margins of safety of the full scale prototype at a particular time which will be called the experiment time. This is to be done in a flight environment sufficiently similar to the prototype environment that the same phenomena are encountered. It is most significant to note that the primary difference between this approach and the traditional one is that this is similitude from a practical engineering viewpoint rather than a completely theoretical viewpoint. The latter approach to the problem usually leads to the conclusion that anything short of one-to-one matching is unacceptable o r at least highly questionable.

3

Exact similitude is impossible because aerodynamic, thermal and structural modeling do not have the same scaling laws. Some physical factors such as thermal contact resistance, boundary layer transition between laminar and turbulent flow, and structural details a re impossible to scale, and the structural concepts being considered for HL-10 are very involved.

Even if perfect similitude (by a one to one correspondence of all parameters) was obtained, the distortions arising in an actual structure caused by attachments, resulting in thermal shorts or leaks and structural edge effects and differences in edge fixity, would negate, in any physical structure of this type, the identical similitude calculated analytically.

However, this in no way negates the utilization of a test bed, such as the SLAMAST vehicle, to obtain valuable and useful information for the HL-10 and other lifting re- entry vehicles of this class. SLAMAST can be used to confirm o r correlate the three basic elements of design; i. e., the environment, the material characteristics, and the analytical and design tools. In particular, it can provide:

additional knowledge about the aerodynamic environment associated with lifting re-entry vehicles.

a means to check the adequacy of present analytical and empirical methods of analysis for pressures, transition points, flow separation, and effectiveness of control surfaces.

additional knowledge about the thermal environment, and the heat transfer to and through the vehicles, including ablation rates.

a means to check the adequacy of present analytical and empirical methods of thermal analysis for convective heat transfer, ablation, conductive through complex structures including honeycomb sandwich, and multilayer insulation.

a means of checking structural methods of analysis and design procedures for complex structures of this type subjected to thermal loads, in-plane loads, and lateral pressure loads.

a means of checking the actual performance of material systems under actual flight conditions, and comparing them with data obtained by laboratory tests, including the ablation and bonding materials.

To insure that valuable information is produced which is applicable to prototype lifting re-entry vehicles:

(1) It is necessary that SLAMAST have a sufficiently similar environment to include al l phenomena that will exist in the flight of a prototype lifting re- entry vehicle.

(2) It is most desirable for the structural test elements of the SLAMAST to use the same materials a s the prototype elements on the full scale vehicle.

4

m II 1 I I 1 1 SI 3

It is necessary for the SLAMAST test components to employ the same structural concepts as the corresponding prototype subsystems.

It is necessary that the test elements represent the same thermostructural response characteristics.

The identical factors of safety used in the design of the prototype should be used in the design of the model test elements.

The SLAMAST test section must be designed to have the same critical mode of failure as the prototype section.

The margin of safety for the critical loading condition of the prototype section should be identical to the margin of safety of the test element under the same loading condition. The h i e e r margins of safety should be reproduced in the same proportions as in the prototype to a degree dependent on their closeness to the minimum margins to accommodate interaction effects.

This final requirement needs some additional discussion since it is an important key to the whole approach. If the minimum margin of safety is 0.01, for example, and the next lowest one is 0.5, again for example only, this clearly indicates that the mode of failure represented by the 0.01 is certainly the weak link with all other potential failure modes fa r removed from possibility. This would be the case unless there are major inaccuracies in the environment, material property, o r structural response predictions. If on the other hand, the minimum margin is 0.01 and the next lowest is 0.015, it could be most important to reproduce b d h these margins in the test panel. The reason for this is that even though the mode of failure represented by the 0.01 margin is predicted to be critical, in reality the mode represented by 0.015 could be the critical one due to tolerances on all factors going into the prediction process. This is shown in Figure 3. It is seen that the tolerances could act in a way to make mode tralr the actual critical mode, whereas mode "b" was predicted to be the weak link. Therefore, the goal in trajectory and test panel design will always be to reproduce, at SLAMAST experiment time, all the HL-10 margins of safety. But since this is an extreme condition to insist upon, it will be sufficient to reproduce only the minimum margin and others sufficiently close to the minimum as to represent potential critical modes of failure.

Another important factor which must be considered in the test design is that the margin of safety for the critical mode in the test panel reaches its minimum value at the experiment time, and that no other margin reaches a value at other flight times which could possibly fai l the test panel prematurely.

This philosophy will derive the greatest value from SLAMAST and insure that the prototype vehicle will accomplish its mission. By this procedure, the SLAMAST will uncover, by suitable instrumentation, any unsatisfactory knowledge of the environment, any deficiencies in analysis, design, or material behavior, and any unforeseen phenomena. In other words, by the SLAMAST flight profile simulating as nearly as

possible the critical aspects of the prototype flight profile, by using the same materials and the same structural concepts, and by forcing the model structural element to have the same critical mode of failure, with the same factors of safety and the same margin of safety, the most important elements of thermostructural simulation a re achieved, without the complications of less important considerations.

The test panel designed with this philosophy will be equally close to failure as the full scale prototype. This is perhaps a unique feature of this approach since many models are designed such that failure thresholds a re fa r removed from expected loading levels.

Even though a large number of the critical conditions for the HL-10 structure occur at approach to touchdown, o r abort, through the use of the philosophy stated above, the SLAMAST can study these critical conditions although it has no abort o r touchdown considerations itself.

Test Panel Design and Trajectory Shaping. - A s discussed above, the goal in the design of the test panel and the shaping of the trajectory is to arrive at a situation, at a specific SLAMAST flight time (the "experiment time"), such that the margins of safety against the various modes of failure are all the same as the margins in the HL-10 or in whatever prototype is being simulated. It is not expected that this can be completely accomplished since this would be perfect similitude, most difficult, if not impossible to achieve. It is necessary only that the minimum one, o r ones be matched. To assist in shaping the trajectory to reach the desired experiment conditions, the classified similarity parameters germain to rectangular panels undergoing small deflection a re used. These parameters deal with buckling, yielding, fracture, thermal and mechanical stress, extensional and flexural stiffness, etc. which are representative of the type structure under investigation.

The procedure for designating the SLAMAST trajectory and panel design to simulate on HL-10 panel design is demonstrated in Figure 4.

It is conceivable that one flight of the SLAMAST vehicle would contain more than one experiment. By carefully designing the trajectory, it is possible that one critical load condition for the prototype can be simulated at one time in the SLAMAST flight while other conditions a re being simulated at other times. The feasibility of achiev- ing this is further enhanced due to the fact that there a re two thermostructural experiment locations on SLAMAST, namely, a windward panel and a leeward panel.

Simulation of the prototype entry environment with the SLAMAST vehicle must include operating the SLAMAST vehicle in an entry corridor that results in heat transfer rates comparable to those anticipated for the prototype. Figure 5 summarizes the peak heat transfer rate distribution predicted for the HL-10 vehicle for the various design trajectories. Superimposed upon Figure 5 is the predicted SLAMAST environment for entry velocities of 20, 000 fps (6.58 km/s) and 25,000 fps (8.22 km/s) and entry path angles of 1 and 10 degrees dfh. Note that the environment of the SLAMAST

6

8 8 I

vehicle is comparable to the HL-10 environment ranges for the complete range of path angles at VE = 20,000 fps (6.58 km/s). However, at an entry velocity of 25,000 fps (8.22 km/s) the SLAMAST environment is well above that of the HL-10 and well outside the efficient operating range of typical low density ablators. Hence, it appears that an entry velocity of 25,000 fps (8.22 km/s) provides much too severe an environment at the steeper path angles. In addition, note that for an entry velocity of 25,000 fps (8.22 km/s), the SLAMAST heat transfer environment falls well above the HL-10 environment, in a regime where most low density ablators exhibit relatively poor performance.

In addition, the effect of guidance e r r o r s must be considered. A nominal toler- ance of * 3 /4O on entry path angle exists. Figure 6 illustrates the effect of entry path angle and velocity on maximum stagnation heating rate. Since the local heating distribution is proportional to the stagnation heating, similar trends with path angle will exist for both the stagnation and local body points. Note that the variation of maximum heat transfer rate with entry path angle is much smaller for the 20, 000 fps (6.58 km/s) entry velocity than for the 25,000 fps (8.22 km/s) case. Hence, guidance e r r o r s would introduce less variation in local heat transfer for the 20,000 fps (6.58 km/s) case than for the 25,000 fps (8.22 km!s) case and thus make the 20,000 fps (6.58 km/s) entry case a more desirable one for experimental purposes.

The environmental simulation technique for trajectory definition requires two major inputs:

(1) Histories of those environmental parameters which must be matched to satisfy a particular experiment.

(2) Closed form expressions for the instantaneous local values of the environmental parameters.

The histories of requirement (1) may be a function of velocity or time. Two or more parameters may be treated simultaneously; however, solution time increases exponent- ially with the number of simultaneous parameters. When multiple parameters are considered, weights reflecting the relative importance of each must be assigned to the parameters. The closed form solutions of requirement (2) may, in general, contain te rms which are tabular functions of other variables. With these inputs plus an initial trajectory to start the iteration, the ODC (Optimum Discrete Control) digital program can be used to generate the SLAMAST trajectory which optimizes the weighted match with the experiment parameters. The comprtation time required to iterate to a solution is a function of the initial trajectory and can be reduced by using parametric studies to provide a judicious initial trajectory profile. One important environmental parameter which is useful in simulation studies is heat rate. Figure 7 shows integrated stagnation heat ra te for the initial SLAMAST trajectory and the eleventh iteration made by the ODC program in attempting to match the desired stagnation heat history indicated. Further iterations could be made to improve the match, if necessary, since the program was still converging on the desired heating history at the eleventh iteration.

7

Other SLAMAST benefits

Boundary layer transition studies: Magnitude of the local heat flux is strongly influenced by the state of the local boundary layer, Le . , laminar o r turbulent. Hence, an optimum lifting entry vehicle design is influenced by the boundary layer transition cri teria employed.

A study of available ground test transition data on SLAMAST-type elliptical cones at an angle of attack w a s undertaken to determine whether existing sphere-cone transition prediction techniques, can be applied to an elliptical configuration. It was con- cluded that because of the paucity of transition data on elliptical bodies, a valid comparison with sphere-cone transition data could not be made. Therefore, no method is currently available by which a logical extension of sphere-cone correlation techniques will allow qualitative estimates of the expected conditions under which transition wi l l occur on an elliptical vehicle at an angle of attack.

An additional application of the SLAMAST vehicle would be to measure the onset and propagation of boundary layer turbulence as it is influenced by the various com- binations of flight parameters like angle of attack, surface temperature, mass injec- tion rate, local Reynolds number, and local Mach number. This understanding of boundary layer transition can then be logically extended to application on the larger prototype vehicles with the subsequent optimization of the low density ablator heat shield.

Materials performance: Over the past several years, ablation performance anomalies have been noted in both low and high density materials. Several of the anomalies are described below:

(1) Surface sealing of silicone elastomers observed in Langley 2500 KW ARC - leads to uncontrolled swelling.

(2) Boundary layer combustion observed on phenolic nylon at Hi oxygen partial pressures.

(3) Boundary layer combustion observed on low density filled epoxy in honeycomb (Avcoat 5026-39-HC G) in low heat flux regime.

(4) Significant surface recession occurs on low density filled epoxy in honeycomb (Avcoat 5026-39-HC G) for low pressure, low heat flux, low aerodynamic shear conditions due to combustion of pyrolysis gases at char surface.

A s more emphasis is placed on optimization of the low density class of ablators, a good understanding of the in-depth chemical reactions as well as the aerothermo- chemical interactions of the injected decomposition products with the hypersonic boundary layer is required. The complex chemical reactions are dependent on several parameters of which temperature and pressure a re of key importance. Internal pressure buildup is directly a function of the molecular weight and rate of the gas generation, and the char layer porosity. The char porosity is very dependent on time; hence, only

8

f I t I I I

a true simulation of a time history of the prototype environment will result in an accurate duplication of the in-depth and surface chemical reactions. SLAMAST will be a n excellent tool to help illuminate these materials behavior phenomena.

Aerodynamic Flight Experiments: SLAMAST class test beds provide excellent facilities for investigation of aerodynamic characteristics in a true re-entry environment not otherwise obtainable in present ground test facilities. The SLAMAST flight investigations wil l provide aerodynamic data of considerable interest in the hypersonic-low-density flight environment over a large portion of the trajectory. Although these data may be of limited scope (depending on the primary experiment), they will include control effectiveness (derived from trim angle of attack as a function of control deflection), associated control loads, control duty cycles, stability characteristics, and lift and drag performance. These data are of importance to l if t and drag performance. These data a r e of importance to lifting vehicle technology and will be correlated with available ground test and analytical results. SLAMAST flights specifically oriented to provide aerodynamic information offer even greater op- portunities for significantly advancing aerodynamic technology.

I

Other testing alternatives. - At present, there a r e only two approaches available to verify or validate a design of a manned lifting re-entry vehicle, or for that matter, any flight system. These a r e flight testing of a prototype vehicle o r individual ground testing of critical items.

Although the actual flight testing of a prototype vehicle under its associated real flight environment does provide the most definitive data, the costs associated with these design verifications are usually extremely high. The second approach, namely ground testing, while reduced in overall cost, does however only provide a limited amount of data on singular phenomena (i. e., the ground test results usually do not include interaction effects, which are quite important in total structural response). There are basically four types of facilities in which structural re-entry environments simulations can be performed. These are: (1) large conventional wind tunnels; (2) a rc heated tunnels, (3) rocket exhaust, or (4) pressurized chamber with radiant heat sources.

Few wind tunnels are available which are large enough to accommodate test panels of the sizes proposed on SLAMAST { about 14 inches (. 35 meters) by 16 inches (. 42 meters) } . None of these facilities can provide experiment times available on a flight vehicle such as SLAMAST.

At the present time, there a r e no arc heated facilities which a r e capable of operating for long times with a large flow field, nor, are there any planned for the near future.

Rocket exhausts such a s those at the Malta facility are generally used for thermostructural tests. However, for the conditions being considered, they a re not

9

acceptable because the maximum flow field is limited to 15 inches (. 38 meters) and the heat fluxes and pressures will be much too high for a lifting vehicle environment.

The remaining choice would be to place the test panel in a chamber designed that a differential pressure could be applied to the panel, with radiant heat sources providing the necessary thermal environment. If this w e r e possible, which it is not today, it would be a rather expensive process.

The facilities described above would not be capable of varying the heat flux o r pressure during testing, which would more closely simulate actual flight conditions. This effect could be minimized if the facilities had enough latitude on environmental control, i. e., such that those environments producing the minimum margins of safety in the critical modes of failure of the test panel could be simultaneously achieved. I t is essential that the thermostructural response of the test panel is not seriously impaired by producing artificial environments either too quickly o r at unacceptable rates. The entire discussion in the philosophy section highlights the goal of matching as closely as possible the similarity parameters during the earlier flight times so as to approach the "critical conditionst' in the same way as in the prototype vehicle.

The real limitations of ground testing lies in its inability to combine simultaneously the effects of aerodynamics, variable heating rates, ablation phenomena, inertia loading and dynamic loading which can only be achieved in flight. It is particularly important when the design margins of safety for several of the most critical failure modes are close. The interaction of these failure modes under a combined environment may seriously affect the thermostructural response of the test specimen.

It is during these combined environments that unforeseen events occur: events that were not predicted on the basis of ground tests, which at best could only combine two parts of the flight environment, and not at the desired level, at that. For example, only after flight testing in which aerothermal effects w e r e included was it found that the fuselage and tail surfaces of the X-15 research airplane had to be reinforced. It w a s determined that the vibration problem exhibited by these surfaces w a s compounded by thermal stresses that were induced simultaneously by other factors.

Flight tests a re also extremely useful in uncovering unrelated but significant phenomena. On the GTV program, which w a s primarily aimed at evaluation of ablative materials characteristics in re-entry environments, the phenomenon of roll resonance w a s observed. This completely unpredicted phenomenon, it was found, can seriously affect the total performance of the re-entry system unless special precautions a re taken. SLAMAST flight tests, aimed at simulating as closely as possible the actual flight conditions and environments of the prototype manned vehicle could well reveal phenomena, heretofore unknown, since the flight regime associated with lifting re-entry vehicles and their associated behaviors are still in the early stages of investigation.

10 1 l

DESIGN FEASIBILITY

This section of the report summarizes the system design effort performed during the SLAMAST study.

It includes a general description, Aerodynamic, Flight Dynamics, Thermodynamics, lnstrumentation and Recovery Studies as well as subsystem trade-off studies, subsystem specifications and preliminary hardware definition of the following subsystems:

(1) Attitude Measurement and Control

(2) Instrumentation and Communication

(3) Recovery

(4) Electrical Power and Distribution

(5) Separation

(6) Structural Design and Packaging

In addition, an Integrated Ground Test Plan and general descriptions of the mechanical and electrical Aerospace Ground Equipment (AGE) are presented along with a program description and detailed schedules

The proposed system is the result of several interations. A s originally conceived, it was thought the re-entry velocities of 25K fps (8 .22K ids) o r greater would be required to obtain environmental similitude with HL-10. This belief necessitated the use of a 4 stage Scout system and required a spacecraft length of approximately 48 inches (1.46 meters), the maximum which could be accommodated within the Scout shroud.

Subsequent progress in the trajectory analysis resulted in an established requirement for re-entry velocities of 20K fps (6.58K m/s) o r less. This allowed the use of a 3 stage Scout booster and relaxed the geometric constraints so that it was not necessary to use a 48-inch (1.46-meter) spacecraft. This, in turn, permitted a design consideration of spacecraft length as a function of approximate historical packaging effi- ciency using non 'f tailor-m ade' components.

Although this was not a design optimization effort, the "test-bed" philosophy dictated a preliminary consideration of the costs involved and a reasonable effort to keep these costs compatible with the technical program objectives. Final selection of the "feasibleff subsystems for design purposes was, therefore, predicated upon econo- mic viability a s well a s technical feasibility.

11

The "test-bed" philosophy also resulted in consideration of the inherent versatility of the as-designed configuration and of alternate component or subsystem selections which could increase specific spacecraft capabilities.

The SLAMAST spacecraft design chosen for feasibility demonstration is a modified elliptic cone vehicle with a length of 63 inches (2.07 meters) and a weight of 330 pounds (150 kilograms). The ellipse major axis is 28.04 inches (0. 805 meter) and the minor axis i s 1 6 . 0 4 inches (0.788 meter).

It incorporates two flaps; the top (or pitch) flap is servomotor powered for variable angle control, whereas the bottom (or drag) flap is driven by a pneumatic one-shot device to a constant angle. These flaps a r e faired into flat areas on the spacecraft. This reduces the flap pivotal point design and thermal insulation problems without significantly affecting the aerodynamic performance of the craft. The vehicle has a nominal W/CLS of 250 lbs/ft2 (1232 Kg/m2) with a maximum lift-to-drag ratio of 2 .4 at 8 degrees angle-of-attack.

The selected structure is aluminum and is comprised of two integrally machined keel members which are stiffened by a ser ies of webs and formers. This configuration provides integral hard points for separation and recovery subsystem attachments and is compatible with a removable access and test panel concept.

A mechanically attached graphitic nose tip and an ESM (GE-Silicone) heat shield a re provided for thermal protection.

Figures 9 and 10 show the general configuration discussed above. Figure 11 is an inboard profile showing the internal packaging arrangement and Figure 12 shows the test panels. Tables I and TI show the weight statement and preliminary component list for the spacecraft. The instrumentation and communication subsystem consists of a C-band tracking and ground command link, an S-band PCM continuous data transmittal link (utilizing a "micro-miniaturized" multicoder) and record and playback capability. It includes those diagnostic and performance sensors necessary to gather the data defined in the measurements list.

The recovery subsystem is a subsonic subsystem and consists of a FIST ribbon decel parachute, a ballute, and a n rf beacon location aid. The ballute is ram-air in- flated and provides inherent spacecraft flotation capability. (The drag flap mentioned above may be considered a part of the retardation sequence prior to parachute deployment. ) This system has been sized to provide a spacecraft water impact velocity of 5 100 fps (32. 9 mps).

The separation subsystem consists of four collet assemblies, an in-flight disconnect (IFD) and auxiliary hardware. The collets a r e "finger and piston" devices which mechanically attach the spacecraft to the spacer. They a re pneumatically operated and, upon command, gas pressure pushes the pistons forward. This releases the mechanical attachment and, by continued piston travel, imparts a separation velocity to the spacecraft.

1 2

The electrical power and distribution subsystem consists primarily of harnesses, a power switching module, a battery, an in-flight disconnect (SC to spacer) and a spacer-to-shroud-access-door umbilical for ground power accommodation.

The selected attitude measurement and control subsystem derives its reference orientation from a Whittaker PRYS platform. It utilizes a programmed sequence for mission accomplishment. Exospheric, 3 axis reaction control and attitude orientation after separation is provided by a Freon-14 gas expulsion system. Upon re-entry, the pitch control (only) is relinquished to the top flap. Roll and yaw control a r e main- tained through the Freon-14 system. The design trajectories were characterized by a constant altitude trajectory from pull-out to drag brake initiation and this is obtained by modulation of the pitch flap angle, roll nulling, and yaw rate limiting during flight in the atmosphere. At the end of the experiment, the drag and pitch flaps are used in conjunction with each other to begin the ballistic mission termination.

The experiment specimens consist of two panels; one on the top (leeward) surface, and another larger one on the bottom (windward) surface. The reference design for these panels w a s based on thermostructural modeling requirements.

A cylindrical spacer is proposed. This spacer provides the spacecraft-to-booster mating interface. It i s designed so that the spacecraft-to-spacer mechanical attachment takes advantage of the inherent hard-points in the spacecraft. The spacer as- sembly also includes the mountingarm*of the umbilical for ground power access and the separation subsystem hardware previously described.

Capability. - A s shown, the SLAMAST spacecraft will endure boost loads and will separate from the third stage of Scout at a timed interval from booster engine cut-off. It will orient itself at a predetermined pitch angle-of-attack (up to a maximum of 16 degrees) and will stabilize itself in roll and yaw within a &5 degrees dead band.

It will re-enter, pull out of its initial ballistic path (by using aerodynamic force), and will re-establish the time reference based on the pull-out maneuver. (This com- pensates for booster burn out altitude variations with resulting e r r o r effects on separation time from launch and its potential perturbation of range, )

The angle-of-attack is modulated (through the programmed p2ich flap control) to fly the predetermined trajectory.

*Modification of the umbilical mounting a rm would permit the selected SLAMAST to be flown on 3 stage SCOUT without the shroud.

13

A f t e r the timed glide portion of flight, the spacecraft will begin its mission termination by deploying both flaps. After a set time from this event, the recovery sequence proper begins with the deployment of the FIST type ribbon parachute.

The spacecraft wil l survive water impact, at 100 fps (32.9 mps), and will subse- quently float for 72 hours minimum. The rf beacon location aid will operate for a minimum of 10 hours.

During boost, re-entry, and glide, all information is gathered from the included sensors and is transmitted continuously over an S-band PCM system. During re- entry and glide, this information is also recorded. During mission termination (both flaps deployed) the continuous data transmittal and recorded data playback occurs simultaneously. and check-out the spacecraft has also been included in this capability.

The Electrical and Mechanical AGE required to handle, test, ship,

Flexibility. - Boost and Re-entry: The boost phase design loads arose from the 3 stage Scout environments as reflected in the Systems Specification, and the proposed design will meet these requirements. ,

The design bounds on re-entry path angle were from yE = -1 degree to yE = -10 degrees. The craft is capable of re-entering with any between these bounds. It is also flexible enough to re-enter at either steeper o r shallower path angles. The factors which would have to be assessed against a specific YE outside the -1 degree to -10 degree bounds are: heat shield, attitude sensing and reaction control impulse requirements .

Pull-Out: The pull-out altitudes specified from the feasibility demonstration point of view are predicated on a pitch flap initial deployment angle of 40 degrees and a re dependent primarily upon yE (i. e. yE = -1 degree causes a pull-out at approximately 145K feet (44.2 Km) and yE = -10 degrees causes a pull-out at approximately 115K feet (35.06 Km) . This attitude could also be varied by changing the initial pitch flap deployment angle (6

equal to 40 degrees are: heat shield, structural strength, and spacecraft stability.

40 degrees for higher pull-out, 6, < 40 degrees for lower pull-out). The factors w c 'ch would have to be assessed against specific 6, not

Glide: The pitch flap program chosen for feasibility demonstration provides for constant altitude flight through modulation of the angle-of-attack. This program could be modified to provide for angle-of-attack/altitude composite variations, etc. ) . The factors which would need to be assessed against a specific glide path control scheme are: pitch flap actuator capability, available electrical power, programmer modifica- tions, heat shield requirements, spacecraft stability.

The roll control implementation scheme chosen for feasibility demonstration provides for roll nulling within a deadband of &5 degrees (from separation through to parachute deployment). This could be varied by establishing a desired roll history program. For example the craft could be banked approximately 90 degrees, kept in

14

that attitude for a specific time then reverse banked through 180 degrees and finally re- turned to null for recovery initiation. The primary factors influenced by implementing something other than roll nulling are: programmer changes and total impulse and thrust level changes. (A significant increase in available impulse and thrust level could be provided by using a hot-gas system. This system also requires significantly less volume fo r a given impulse than the Freon-14 system).

The level flight duration chosen for design feasibility i s approximately 500 seconds. This can be varied by a simple timing change. The factors which would need to be assessed against a specific gl !(le duration are: heat shield (thermal insulation) requirements, retardation effectiveness for recovery, and reaction control impulse requirements.

Recovery: A l l design trajectories employed the drag flap a s a retardation device and showed subsonic, low dynamic pressure conditions at parachute deployment. If less high altitude retardation was desired for higher ballistic termination velocities, the subsonic recovery system could be elevated to a Mach 1.5 capability by changing to a hemisflo parachute with small weight and volume penalties.

Control System: The attitude reference system chosen for design feasibility is a modified MARS platform (termed PRYS by Whittaker Corp.). The mission accuracy predicated on the e r r o r s derived from this platform (and ancillary hardware) is presented in Section 4.1 of Volume II. Other reference systems are available which could, at some increase in expense, increase the total mission accuracy.

Redundancy. -

Data Transmittal: The design provides for both continuous real time S- band data transmittal and data recording and playback without interruption of the real time capability. This assures data gathering even if factors such as plasma attenuation o r ground receiver availability prevent continuous reception of telemetered data.

The playback mode u7ill be initiated by either of two events:

(1) Primary, the passage of a specified time from pull-out.

(2) Back-up, the receipt of a ground command through on-board C-band via the FPS-16 radar.

Mission Termination: The deployment of the drag/pitch flaps for mission termination initiation will be effected through either of two events:

(1) Primary, the passage of a specified time from pull-out.

(2) Back-up, the receipt of a ground command through on-board C-band via the FPS-16 radar.

15

Structure: The experiment panels are mounted so that they form the "lid" of a structural pan. The pan itself is an integral part of the spacecraft structure and has its own heat shield and structural integrity. A failure of the panel should not result in spacecraft failure.

Destruct: Although no design implementation of a destruct system was pursued, the wealth of reference sensors and performance data measurements assure the availability of existing on-board equipment which can be used to signal the need for de- struction and to provide initiating action to the chosen devices. The sensing of this need would be derived from totally redundant parameters.

Pyrotechnic Initiation Devices: All pyrotechnic devices will be provided with redundant and isolated circuits.

PROGRAM CONSIDERATIONS

This section of the report describes the program devised to translate the SLAMAST spacecraft design developed during the design feasibility study into flight hardware. The cornerstone for this description is the overall program schedule, the detailed analyses schedule, subsystem development schedules and test schedules (See Figures 13 through 16).

The program philosophy adopted for development of the SLAMAST spacecraft was largely dictated by the premise that SLAMAST would be a test bed for conducting a variety of experiments. A s a result, emphasis was placed on maximizing confidence in the test bed developed in order to maximize the opportunity for conducting the experiments. This philosophy had its major impact in two areas . The first was in the design definition and the second was in the development schedules and tasks.

In the design definition, the emphasis was placed on simplicity and specification of conventional subsystems and components in order to maximize confidence. For example, a relatively simple preprogrammed flight path control scheme was adopted; a conventional cold gas reaction control subsystem was specified in lieu of a less conventional hot gas subsystem which had lower weight and volume requirements; the basic attitude measurement sensor selected was one which is relatively inexpensive and has a substantial flight history; the recovery subsystem utilizes a conventional subsonic decel chute and location aids along with a somewhat less conventional but flight proven ballute; the structural design is a conventional semimonocoque design utili zing conventional materials and fabrication techniques ; s tructur a1 de sign margins of safety were kept high; the separation subsystem was based on designs which have been successfully fabricated and flown on other GE-RSD programs. Thus, with the exception of the S-band communication system, which was a contractual requirement, the spacecraft could be characterized as being composed of yesterday's subsystem designs. Even in the S-band communications area, hardware specification was based

t I 8 1 1

8 8 t 8 8 I t 8

on the current state of development and well defined plans for development so that major breakthroughs in the state-of-the-art are not required for successful implementation.

The program schedules developed were based on the assumption that the design evolved during this study would be implemented, and that a conventional four stage information release system would be implemented. The data generated during the study satisfies the requirements for a stage 1 release. A s such only three more stage releases will be required during the development cycle. Figure 13 shows the overall program schedule that was evolved. The schedule shows only the elapsed time from go-ahead to first flight since this is a key input to a project development plan and sufficient data exists to develop this schedule. Subsequent flight schedules would be dependent upon definition of the experiments to be conducted, and the philosophy of their conduct, e. g. are the tests to be independent of one another or a r e subsequent test definitions going to be based on the results of previous tes ts?

The analysis schedules a re shown in Figure 14, the subsystems development schedules are shown in Figure 15 and an integrated test schedule is shown in Figure 16. In general, stage 2 information releases (PDR) a r e coordinated to occur 3 months from go-ahead and stage 3 releases (FDR) are scheduled to be made approximately 10 months from go-ahead. Note that completion of order of qual hardware is always scheduled to occur after fourth stage component releases are made. This is in keeping with the conservative approach adopted. Frequently, on other programs, qual hardware is ordered on the basis of third stage release data.

It should also be noted that the program defined includes an engineering development vehicle (EDM) and a qual vehicle (prototype spacecraft), a s well as a flight vehicle. It is possible, and has been done on other programs, that the development vehicle or the qual vehicle can be eliminated if the program philosophy is changed. Another element of the schedule that can be subject to revision is the adopted philosophy that component qual will be completed before system qual tests are initiated. Pro- grams have been conducted where the requirement was to achieve component qual only some time before flight. Finally it should be noted that before the qual spacecraft is subjected to the qual environments, it will be subjected to acceptance test environments. On other programs, this s tep is frequently omitted.

In summary, the spacecraft development schedule evolved results in a flight two years after go-ahead. Assuming a program start in mid 1968, this means the first flight would occur in mid 1970. This time span is believed to be in keeping with a relatively conservative program philosophy. A shorter schedule with potential eco- nomic benefits can be generated by implementing any o r all of the alternatives iden- tified in previous paragraphs.

17

TABLE I. - SLAMAST WEIGHT & BALANCE

18

ITEM

(l)Heat Protection

E le c tri c a1 Equipment

T/M & Tracking

Attitude Contr. & Meas.

Recovery

Flap Control

Brake Control

Re ac tioii Control

Nose Tip/Ballast

rota1 Spacecraft

Adapter Structure

Separation System

rota1 Adapter

rota1 S/C + Adapter

~~

45.1

88.7

19.0

27.0

22.7

11,8

7 . 0

3.0

36.9

68.8

330.0

22.4

4.9

27.3

357.3

NOTES - (1) Includes Flaps & Test Panels

C.G. STA. (IN. FROM NOSE)

46.2

40.1

37.2

41.0

26.7

59.0

59.0

59.0

53. G

7 . 1

35.8

81.2

78.0

80.6

39.2

MOMENT OF INERTIA

(SLUGFT~)

IRoll = 2.24

IYaw = 28.7

IPitch = 27.4

I - \ The weights etc. stated above are the result of the last iteration and have not been totally factored into the report. However, the differcnces arc not crucial to the feasibility demonstration.

I 1 I 1 I tl 8 1 t

I 1

TABLE II. - SLAMAST: PRELIAIINARY PARTS LIST

~~

1. Recovery

2. TT&C W C )

3. Separation

4. AM&C

5. EP&D

Component

1.

2.

3.

4.

5.

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

11.

1.

2.

3.

4.

5.

6.

7.

1.

2.

3.

4.

5.

6.

7.

1.

2.

Drogue chute Assy. . (Campy. risens, and deployment bag)

Ballute Assy. . (Ballute and deployment bag).

R-F b e a c ~ n & antenna Drogue Chute Release Mech.

Ejection Mortar & Squib

Telemetry Antenna

Telemetry Transmitter (S-bat

PCM Multicoder (Micromin)

Baseband Assy Unit

Tape Recorder

Analog Signal Conditioner

Sensor Power Supply

C-band Transponder

Decoder Module

Single A x i s Accelerometer

Sensors: Pressure Temp etc.

Collet Assy.

Freon Tank, 3000 psi

Pyro. AduatedValve

Fill & Vent Valve

Pressure Transducer

IFD - 26 pin

Separation Switch

Autopilot

PRYS Attitude Ref. Unit

Tri-axial Rate Gyro

Tri-axial Accelerometer

Controller Programmer

Pitch Flap Actuator

Reaction Control

a) Tanks; 3500 psi

b) Pressure Regulator

c) Filter

d) Solenoid Valves

e) Pyro Actuated Valve

fi Nozzles

Battery

Power Controller

1

1

1

1

1

1

1

1

1

1

1

1

1

1

3

22 DO

4

1

1

1

1

1

1

1

1

1

1

1

1

6

1

1

6

1

6

1

comments

(Goodyear, Aerospace)

(Goodyear, Aerospace)

Similar to Microcom Model T-40

similar to MK-12 (recent developmental award)

Similar to Borg-Warner R310

Similar to GE/RSD 47C138575

Similar to GE/RSD 111C5250

Similar to VEGA Model No. 302 C-2

Similar to VEGA Unit

Similar to Honeywell GG 322

Holex Inc. 509OA

MS 28889-1

Cannon P N 1020570001 (26pin)

GE PN R 2336

Similar to Whittaker MARS GE/RSD 47~142658

GE/RSD 47C142659

Bendix P N 122 dwg 3170164

Similar to Sterer Mfg. PN 99643

GE/RSD 47C139969

Carleton P N 1809001

Holex Inc. 5090A

Yarnley 7AH

19

20

I

I

I? IIYSIC A L MECHANCAL THERRLIAL

I 1

I I LOADS A e1-0 dyn anii cs Inerti a1

STRESS Tensile Shear

HE A TI NG DEFLECTION D Y NA 31 ICs

STRUCTURAI, DESIGN e--l I PREDICTED RESPOhTF: I

Figure 1. - Elements of Structural Design

5 TIME

Figure 2. - Margin of Safety as a Function of Time

a

TOLERANCERAND ON MODE "af1

OSSIBLE OVERLAP

TOLERANCEBAND ON MODE "b"

TIME

Figure 3. - Margin of Safety - Critical Failure Modes

21

t 2 1-391

t

I I

1 t 1

3

Frc 0 d

I

4

I I I

U I 1 22

cv E \ m c-c c-c c 3 c

E s

w 5 k u w

x' c E

ENTRY ANGLE, DEG. ENTRY VELOCITY k\ WW105 ?,ow

20,000

25,000

20,000

25,000

20,000

20,000 - 25,000

H L l O (UNDERSHOOT) \ /

Km/SEC

8.22

6.58

8.22

6.58

a. 22

6.58

6.58

8.22

/ - - H G l O (OVERSHOOT)

- 1 -4 -- --1

- WW - SLAMAST WINDWARD RAY S - SLAMAST SIDE RAY

-

I I I I I I

X/L

Figure 5. - Comparison of Maximum Convective HeatFlux Rates for HL - 10 and SLAMAST Vehicles

23

h) t-c Frc

0 0 0 0 N

0 0 0 l-l

0 0 LD

38s zLJ/nLB ‘ZLVH HZJSNVHL LVZH

I I 1 I I 0 0 0 0 0 0

0 m 0

h) 0 d

In

24

8 8 1 8 I 8 8 8 8 8 I I 1 I I I 8 I 1

a.

7.0-

6.0-

5 . 0 -

4.

3.

2.

1.

o-

@a b Lrr

3 b

. a w" b e p: b s

% x c -

C q 0 b

w b e p: 13 w t.c

a, n

c- E

c -

0 I I I I I 22,000 23.000 24,000

VELOCITY, FPS

Figure 7. - SLAMAST Integrated Stagnation History versus Velocity

25

26

8 I 8 I I 8 8 8 8 1 1 8 I I 8 I I I I

8 i 8 8 1 8 8 8 8 B 1 1 8 8 1 8 8 1 I

Figure 9. - SLAMAST S/C

27

28

I 8

30

I

I I

I

‘

I I I I 1 1 1 I

I i

f - a

1 1

f-i-

31

I

I I

32

I 8 I 8 1 8 I 1 1 1 1 1 1 1 I I I

1 a

w 4 z E

M

3

4

w 2

-m-

-8 E D

-$- -8-

I -

m >

t

1

33

h N I (D

111 N

0 N I N N

3 N

0 N

! i 2

I ! I

I ' I

I D I I D l

4

D

kl hD D

&D

D 0

D

h Y

% 4

A

c

fi

B 2

e a

a D

I I I

1 I

0 D D D

D

@ D

Y

D

D

I"

35

N ~

P I

n v a

I I I I I I

I

6 I I

6 4

t)

6

F L

m

" Q)

D -4

Frc

37

TASK NAME

TLM tracking & command

s/s plan

S/S specifications

Order development hardware

Component development

S/S develapment tests

Component specifications

Order qual hardware

Component qual tests ~

Attitude measurement & control (AMq

s/s plan

S/S specifications



S/S development tests


Order qual hardware

Component qual tests

Recovery

S/S plan

S/S specification

Subcontract effort

S / S development tests


Order qual hardware


Separation

s/s plan

S/S specifications



S/’S development tests


Order qual hardware


Electrical power & distribution (EP&D)

( s/s plan

S/S specifications

System schematic


Elect. system development test

Harness design

Order qual hardware


MONTHS FROM GO-AHEAD

A start 0 Complete

v StageRelease 0

Figure 15. - Subsystem Development Schedule

I I 38

~

0 c- oa rl

W I

h 3

I

2 .rl Y m Q)

Q)

Y

H P

B

0 0 (d

0) 6 Y

U m Q)

E-c

I I

40

preliminary design study of - nasa cr-66399 volume 111 summary 14 april 1967 preliminary design...

Documents